WO2007115446A1 - Compresseur à écoulement radial en cascade - Google Patents

Compresseur à écoulement radial en cascade Download PDF

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Publication number
WO2007115446A1
WO2007115446A1 PCT/CN2006/002082 CN2006002082W WO2007115446A1 WO 2007115446 A1 WO2007115446 A1 WO 2007115446A1 CN 2006002082 W CN2006002082 W CN 2006002082W WO 2007115446 A1 WO2007115446 A1 WO 2007115446A1
Authority
WO
WIPO (PCT)
Prior art keywords
stationary
cascade
blades
blade
moving
Prior art date
Application number
PCT/CN2006/002082
Other languages
English (en)
Chinese (zh)
Inventor
Changzhe Liu
Original Assignee
Changzhe Liu
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Changzhe Liu filed Critical Changzhe Liu
Publication of WO2007115446A1 publication Critical patent/WO2007115446A1/fr

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D17/00—Radial-flow pumps, e.g. centrifugal pumps; Helico-centrifugal pumps
    • F04D17/08—Centrifugal pumps
    • F04D17/10—Centrifugal pumps for compressing or evacuating
    • F04D17/12—Multi-stage pumps
    • F04D17/127—Multi-stage pumps with radially spaced stages, e.g. for contrarotating type
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
    • F02C3/045—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor having compressor and turbine passages in a single rotor-module
    • F02C3/05—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor having compressor and turbine passages in a single rotor-module the compressor and the turbine being of the radial flow type

Definitions

  • the present invention relates to the field of aviation turbojet engines, and more particularly to a radial cascade compressor for an aerial turbojet engine.
  • BACKGROUND OF THE INVENTION Compressor is one of the important components of an aviation turbojet engine. Its function is to compress the airflow entering the engine. The compression process of the intake air is a necessary process for the thermodynamic cycle of the aeroengine. The size of the compression ratio is the most important factor determining the engine power. element.
  • the existing small air jet engines are designed with a centrifugal compressor. However, the centrifugal compressor is only compressed by the free-rotating gas centrifugal force. The compression ratio is small, the efficiency is low, and the required compression ratio is not achieved.
  • the compressor is barely suitable for use in engines in avionic generators.
  • Another axial compressor which is commonly used in existing jet engines, is inconsistent with the direction of compression of the airflow due to its centrifugal force, and the centrifugal force can interfere with compression.
  • the axial cascade compressor cannot be made small. Dimensional, small-sized blades are difficult to machine and install, so axial-blade compressors are only suitable for large and medium-sized jet engines and large aircraft. At present, small aircraft are also driven by propeller engines.
  • SUMMARY OF THE INVENTION The technical problem to be solved by the present invention is to provide a radial cascade compressor which has a high pressure and a reduction ratio and is applicable to a flat aircraft engine and a small aircraft turbojet engine.
  • the technical solution adopted by the present invention is: providing a radial flow cascade compressor including an air inlet, a casing, a bucket disk and a stationary blade disk at an axial center; the moving and static The leaf discs are respectively provided with a plurality of moving blades and static vanes arranged in a concentric annular shape, and the moving and stationary vanes respectively form an annular moving and stationary vane grid; the moving and stationary vanes are oppositely mounted, and the circular moving and stationary vanes are mutually connected. Interval; when the bucket disk rotates, the gas is compressed by the air inlet and compressed outward in the radial direction, and then confirmed by the channel row. Out.
  • the bucket disk is a disk having a tapered outer circumference
  • the rotor blade is located on an inner surface of an outer circular region thereof.
  • the moving blade located inside the conical surface of the bucket disc and the relatively stationary vane are both two-dimensional curved surfaces, and the busbars thereof have their height directions parallel to the rotating shaft.
  • the blade area of the inner ring of the bucket disk and the vane disk is larger than the blade area of the outer ring.
  • the blade pitch of the outer ring of the bucket disk and the vane disk is smaller than the blade pitch of the inner ring, and the number of blades of the outer ring is more than the number of blades of the inner ring.
  • the runoff cascade The compressor may also include an axial compression air impeller located within the intake port and mounted coaxially with the bucket disk.
  • annular stationary blade cascade may be provided, the blade shape of the annular stationary blade cascade being a space spiral surface.
  • a non-contact labyrinth seal structure for preventing pressure backflow is provided between the back of the bucket disc and the casing.
  • the radial flow cascade compressor of the present invention utilizes the principle of an axial flow cascade compressor to transform the pipeline layout of the axial flow cascade into a radial concentric annular cascade.
  • Disc construction Since the circumference of the outer ring of the moving and stationary blade is longer than the inner ring, the outer ring can be arranged with more blades; and, because the outer ring has a higher speed, the compression effect with the stationary blade is better than that of the axial flow.
  • Grid compressor Compared with the centrifugal compressor and the axial cascade compressor, the radial cascade compressor of the present invention is suitable for mold manufacturing due to high compression and large design variation range, and can ensure manufacturing precision and low process cost.
  • FIG. 1 is a cross-sectional view showing an assembled structure of a preferred embodiment of the present invention in conjunction with the drawings and specific embodiments.
  • Figure 2 is an exploded perspective view of the embodiment of Figure 1.
  • Figure 3 is a projection view of the position of the moving and stationary blade assembly in the embodiment shown in Figure 1.
  • the radial cascade compressor mainly includes a centrally located air inlet 1, a stationary blade disk 2, a rotor blade 5, and a rear cover. 8.
  • the vane disk 2 has a disk shape, and is provided with a plurality of stator blades 12 and 14 arranged in a concentric annular shape.
  • the stator blades 12 and 14 form an annular stationary blade cascade adjacent to the stationary blade disk 2
  • an annular stationary blade cascade 15 is also provided, and the blade shape of the annular stationary blade cascade 15 is a spatial spiral surface.
  • the moving blade disc 5 is provided with a plurality of moving blades 9 and 13 arranged in a concentric annular shape, and the moving blades 9 and 13 form an annular moving blade cascade; the outer circular area of the moving blade 5 is a conical surface, where The moving blades 9, 13 are located on the inner surface of the conical surface, and the tapered surface disc design prevents the centrifugal force generated by the high-speed rotating moving blades 9, 13 from deforming the vane disc 5 to ensure the stability of the rotating gap.
  • the vane disc 5 is fastened to the shaft 6 and mounted on the rear cover 8 via a bearing 7, the vane disc 2 also serving as a front cover of the compressor, which is fastened to the rear cover 8 to form cabinet.
  • the moving and stationary vanes 5, 2 are oppositely mounted, and the assembled annular moving and stationary cascades are spaced apart from each other and have a reasonable gap to ensure that all the moving and stationary vanes 9, 13, 12, 14 when the vane disc 5 rotates No interference will occur.
  • a non-contact labyrinth seal configuration 4 is provided between the back of the bucket 5 and the rear cover 8 to prevent the compressed flow from escaping along the slit.
  • the annular collecting chamber 11 is located at the periphery of the annular stationary blade grid, and is formed by sealing connection between the stationary vane 2 and the rear cover 8.
  • the compressed gas is accumulated in the annular collecting chamber 11 and then passes through the passage 3. 10 is pressed into the combustion chamber (not shown).
  • the axial compressed air impeller 16 is located in the air inlet 1 and is fastened to the moving blade 5 Tighten the front end of the connected shaft 6.
  • the vane 5 and the vane disc 2 of the present invention The blade area of the inner ring is larger than the blade area of the outer ring.
  • the outer circumference of the moving and stationary vanes 5, .2 is larger than the inner circumference, more vanes can be arranged, and the blade pitch of the outer vane 5 of the present invention and the outer ring of the vane disc 2 is smaller than the inner circumference.
  • the blade pitch of the ring, and the number of blades of the outer ring is greater than the number of blades of the inner ring.
  • the ratio of the number of outer ring blades of the present invention to the number of inner ring blades is much larger than that of the shaft.
  • the height direction of the moving and stationary vanes 9, 13 3, 1 2, 1 4 of the present invention is parallel to the rotating shaft 6, and the moving vanes 9, 13 and the relatively stationary vanes 12 on the inner side of the conical surface of the vane disc 5 are provided.
  • 1 4 are all two-dimensional curved surfaces, and the busbars are parallel to the rotating shaft 6, which is beneficial to the use of the mold to manufacture the parts while fully satisfying the requirements of the airflow compression profile, so that compared with the complex process of the three-dimensional curved blades of the centrifugal compressor
  • the invention greatly simplifies the process, thereby better ensuring product accuracy while reducing the total cost of the engine.
  • the height direction of the moving and stationary vanes 9, 13 3, 1 2, and 14 of the present invention is parallel to the rotating shaft 6, and the moving vanes 9, 13 and the relatively stationary vane 12 on the inner side of the conical surface of the vane disc 5 are provided.
  • 1 4 are all two-dimensional curved surfaces, and the busbars are parallel to the rotating shaft 6 to facilitate the use of the mold to manufacture the parts while fully satisfying the requirements of the airflow compression profile.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

La présente invention concerne un compresseur à écoulement radial en cascade pour un turboréacteur aéronautique. Le compresseur comporte un orifice d'admission d'air (1) prévu au niveau de centre d'un axe, un carter, un disque de turbine mobile (5), un disque de turbine stationnaire (2); une pluralité d'aubes mobiles (9, 13) et des aubes stationnaires (12, 14) en forme de cercles concentriques situées sur le disque de turbine mobile et le disque de turbine stationnaire, respectivement, les aubes mobiles et les aubes stationnaires (3, 13, 12, 14) formant une cascade mobile et une cascade stationnaire, respectivement; le disque de turbine mobile (5) est en regard du disque de turbine stationnaire (2), la cascade circulaire mobile est espacée par rapport à ladite cascade stationnaire; lors de la rotation du disque de turbine mobile (5), l'air est aspiré à travers l'orifice d'admission d'air (1) et l'air comprimé est évacué vers l'extérieur à travers un passage d'évacuation (3, 10). Grâce à l'utilisation du principe de compresseur axial, l'agencement en forme de conduit de la cascade axiale est transformé en un agencement en forme de disque de cascade sous la forme de cercles concentriques radiaux. Selon la présente invention, d'autres aubes sont prévues sur les cercles extérieurs des disques mobiles et stationnaires, et la vitesse des cercles extérieurs est supérieure, l'effet de compression produit par l'aube mobile, en coopération avec les aubes stationnaires est supérieur au compresseur à écoulement axial en cascade. Le compresseur à écoulement axial en cascade selon la présente invention a un taux de compression supérieur et une plage plus large de zone de conception, il est fabriqué de manière appropriée par un moule pour en assurer la précision de fabrication et un faible coût de production, et il est avantageusement utilisé pour un moteur aéronautique et un turboréacteur pour un petit aéronef.
PCT/CN2006/002082 2006-04-07 2006-08-16 Compresseur à écoulement radial en cascade WO2007115446A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
CN200610060217.7 2006-04-07
CNA2006100602177A CN1847664A (zh) 2006-04-07 2006-04-07 径流叶栅压气机

Publications (1)

Publication Number Publication Date
WO2007115446A1 true WO2007115446A1 (fr) 2007-10-18

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Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/CN2006/002082 WO2007115446A1 (fr) 2006-04-07 2006-08-16 Compresseur à écoulement radial en cascade

Country Status (2)

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CN (1) CN1847664A (fr)
WO (1) WO2007115446A1 (fr)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102010047612A1 (de) * 2010-10-07 2012-04-12 Voith Patent Gmbh Laufrad als Verdichter- oder Turbinenrad

Families Citing this family (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN101255873B (zh) * 2008-02-28 2010-06-09 大连海事大学 压气机动叶叶尖小翼
CN102434215B (zh) * 2011-11-07 2014-07-30 王政玉 一种外转子流体动力机
CN102608914B (zh) * 2011-12-22 2014-03-12 西安交通大学 径流式液力透平优化设计方法
WO2013091222A1 (fr) * 2011-12-22 2013-06-27 西安交通大学 Procédé de conception optimisé pour turbine hydraulique du type à flux radial
CN103807022B (zh) * 2014-02-17 2016-04-27 汪一平 一种双向涡叶内冷负压航空氢动力发动机
CN104948300B (zh) * 2015-06-17 2017-04-05 上海理工大学 燃气轮机
CN105201650A (zh) * 2015-10-12 2015-12-30 常胜 一种离心式涡轮发动机

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3941499A (en) * 1974-11-06 1976-03-02 United Turbine Ab & Co., Kommanditbolag Compressor having two or more stages
GB2027811A (en) * 1978-08-09 1980-02-27 Mtu Muenchen Gmbh A gas turbine engine having means for bleeding compressor air
US5025629A (en) * 1989-03-20 1991-06-25 Woollenweber William E High pressure ratio turbocharger
CN1071738A (zh) * 1991-10-21 1993-05-05 株式会社日立制作所 离心压缩机
JPH10231798A (ja) * 1997-02-17 1998-09-02 Kawasaki Heavy Ind Ltd 遠心形多翼送風機
US6085527A (en) * 1997-05-15 2000-07-11 Turbodyne Systems, Inc. Magnet assemblies for motor-assisted turbochargers
CN1616833A (zh) * 2003-11-11 2005-05-18 台达电子工业股份有限公司 含有静叶的离心扇及其扇框

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3941499A (en) * 1974-11-06 1976-03-02 United Turbine Ab & Co., Kommanditbolag Compressor having two or more stages
GB2027811A (en) * 1978-08-09 1980-02-27 Mtu Muenchen Gmbh A gas turbine engine having means for bleeding compressor air
US5025629A (en) * 1989-03-20 1991-06-25 Woollenweber William E High pressure ratio turbocharger
CN1071738A (zh) * 1991-10-21 1993-05-05 株式会社日立制作所 离心压缩机
JPH10231798A (ja) * 1997-02-17 1998-09-02 Kawasaki Heavy Ind Ltd 遠心形多翼送風機
US6085527A (en) * 1997-05-15 2000-07-11 Turbodyne Systems, Inc. Magnet assemblies for motor-assisted turbochargers
CN1616833A (zh) * 2003-11-11 2005-05-18 台达电子工业股份有限公司 含有静叶的离心扇及其扇框

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102010047612A1 (de) * 2010-10-07 2012-04-12 Voith Patent Gmbh Laufrad als Verdichter- oder Turbinenrad

Also Published As

Publication number Publication date
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